The fuselage, wings, and empennage of an aircraft typically include stringers that are coupled to skin that forms the smooth aerodynamic outer surfaces of the fuselage, wings, and empennage. The stringers and skin cooperate to provide flexural and torsional stiffness to these sections of the aircraft. Traditionally, the fuselage, wings, and empennage surfaces and the associated stringers are fabricated from metal, such as aluminum, steel, or titanium.
Fiber reinforced composite materials are widely used in a variety of commercial and military aircraft products as a substitute for metals, particularly in applications where relatively low weight and high mechanical strength are desired. The material is generally comprised of a network of reinforcing fibers that are arranged in layers or plies. The layers include a resin matrix that substantially wets the reinforcing fibers and that is cured to form an intimate bond between the resin and the reinforcing fibers. The composite material may be formed into a structural component by a variety of known forming methods, such as extrusion, vacuum bagging, autoclaving, and/or the like.
As the skins and stringers for various sections of aircraft transition from metallic materials to fiber reinforced composite materials, multiple issues have arisen. In a current fabrication process known as Pre-cure/Co-bond or Co-bond, a fiber reinforced composite skin is formed by stacking layers together that contain reinforcing fibers in a resin matrix. Typically, some of the layers are staggered relative to each other so that the stack conforms to a desired contoured or tapered geometry. The stacked layers are heated and pressurized to cure the polymeric resin matrix and form a precured skin. A stringer mandrel is placed on the stacked layers and stringer plies are stacked onto the stringer mandrel. Pressure and heat are applied to cure the stringer preform using the tooling to form a fiber reinforced composite stringer that is attached to the precured skin. Unfortunately, defects often occur along the interface between the precured skin and the fiber reinforced composite stringer. In particular, the skin typically has a contoured outer surface that includes small steps or drop-offs that are formed by the gaps between mandrel elements of the stringer mandrel. Large gaps cause under compressed areas, e.g., voids, and over compressed areas, e.g., resin poor areas, within the composite stringer. These under and over compressed areas can reduce the load transfer efficacy between the precured skin and the fiber reinforced composite stringer, thereby reducing the rigidity and support provided by the fiber reinforced composite stringer.
The stringer is then Co-cured over the pre-cured skin in a second curing cycle. Pre-curing provides good skin quality due to uniform bagging pressure and relatively simple geometry, but requires two runs of the autoclave process and preparation. Requiring two runs becomes very expensive and time consuming, and introduces risk associated with re-heating the skin laminate.
Another current fabrication method is Co-cure, which allows creating a complex part in a single autoclave curing step. The disadvantage of Co-cure is that bagging over complex shapes creates a non-uniform pressure where the laminate quality can suffer by having resin rich/starved areas, and potential ply distortion due to uneven pressure.
A third current fabrication method is Secondary Bonding, in which all of the constituent components are pre-cured, and then assembled in a bond mold with adhesive and cured together. Secondary Bonding provides the best laminate quality for all parts, but requires multiple cure cycles, increased of preparation, extra tooling, and increased difficulty with fit-up issues of pre-cured parts.
Conventional mandrels include foam mandrels and elastomeric mandrels. Foam mandrels are difficult to remove from the fiber reinforced composite stringer after curing and are often considered “flyaway” tooling that remains in the composite wing. Elastomeric mandrels are easier to remove from the cured composite stringer, but do not offer the stiffness and rigidity required for well-defined edges in the composite stringer. Furthermore, the elastomeric mandrels are typically only available for several uses before the elastomeric material begins breaking down and is not able to withstand the curing process.
Accordingly, it is desirable to provide composite tooling systems and methods of manufacturing composite stringers that are removable from the stringer after curing, are durable, and provide the rigidity to produce well-defined edges on the cured composite stringer. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description and the appended claims, taken in conjunction with the accompanying drawings and this background.